Ring-shaped compliant support

ABSTRACT

A ring-shaped compliant support for a gas turbine engine includes, among other things, an annular case, and an adjustment member that will turn relative to the annular case if exposed to thermal energy.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a divisional of U.S. patent application Ser. No.14/773,190, which was filed on 4 Sep. 2015, which is a 371 NationalStage Entry of International Application No. PCT/US2014/16761, which wasfiled 18 Feb. 2014, which claims priority to U.S. ProvisionalApplication No. 61/774,637 filed 8 Mar. 2013.

BACKGROUND

This disclosure relates to supporting components of a gas turbineengine.

Gas turbine engines typically include a compressor section, a combustorsection, and a turbine section. During operation, air is pressurized inthe compressor section and is mixed with fuel and burned in thecombustor section to generate hot combustion gases. The hot combustiongases are communicated through the turbine section, which extractsenergy from the hot combustion gases to power the compressor section andother gas turbine engine loads.

The turbine section of a gas turbine engine typically includealternating rows of rotating blades and stationary vanes. The turbineblades rotate and extract energy from the hot combustion gases that arecommunicated through the gas turbine engine. The turbine vanes preparethe airflow for the next set of blades. The vanes extend from platformsthat may be contoured to manipulate flow. An outer casing of an enginestatic structure may include one or more blade outer air seals (BOAS)that sit outwardly of the blades and provide an outer radial flow pathboundary for the hot combustion gases.

Vanes, BOAS, and other components are typically found at locationscircumferentially surrounding a center axis of the gas turbine engines.The thermal energy level at a given axial position can vary acrosscircumferential positions. Thus, one circumferential location may behotter than others, due to any number of factors in the manufacturing ofthe gas turbine engine. The blades rotate across the entirecircumferential extent, and thus each experience hotter and less hotareas. The BOAS and vanes, however, are generally stationary and thussome may sit in the hotter positions. This may lead to uneven wear ofvanes, BOAS, and other gas turbine engine components.

SUMMARY

A ring-shaped compliant support for a gas turbine engine according to anexemplary aspect of the present disclosure includes, among other things,an annular case, and an adjustment member that will turn relative to theannular case if exposed to thermal energy.

In a further non-limiting embodiment of the foregoing ring-shapedcompliant support, the annular case defines a bore and the adjustmentmember is received within the bore.

In a further non-limiting embodiment of any of the foregoing ring-shapedcompliant supports, the adjustment member comprises a plurality of wingsextending radially and circumferentially from a ring.

In a further non-limiting embodiment of any of the foregoing ring-shapedcompliant supports, slots in the adjustment member separate theplurality of wings from each other.

In a further non-limiting embodiment of any of the foregoing ring-shapedcompliant supports, the plurality of wings comprises a first array ofwings extending radially and circumferentially from the ring, and asecond array of wings extending radially and circumferentially from thering.

In a further non-limiting embodiment of any of the foregoing ring-shapedcompliant supports, the first array of wings are axially sequentiallyaligned with the second array of wings.

In a further non-limiting embodiment of any of the foregoing ring-shapedcompliant supports, the wings in the first array are axially thinnerthan the wings in the second array.

In a further non-limiting embodiment of any of the foregoing ring-shapedcompliant supports, the wings in the first array are upstream from thewings in the second array.

In a further non-limiting embodiment of any of the foregoing ring-shapedcompliant supports, the wings in the first array are relativelyinsulated and the wings in the second array are relatively uninsulated.

In a further non-limiting embodiment of any of the foregoing ring-shapedcompliant supports, the wings in the first array are upstream from thewings in the second array.

In a further non-limiting embodiment of any of the foregoing ring-shapedcompliant supports, the wings in the first array are axially thinnerthan the wings in the second array.

In a further non-limiting embodiment of any of the foregoing ring-shapedcompliant supports, the wings elongate in response to thermalvariations.

In a further non-limiting embodiment of any of the foregoing ring-shapedcompliant supports, the adjustment member and the annular case havedifferent rates of thermal expansion, such that the adjustment memberexpands more quickly than the annular case in response to thermalenergy.

A ring-shaped compliant support according to another exemplary aspect ofthe present disclosure includes, among other things, an annular case, anadjustment member, and a gas turbine engine component and configured toturn with the adjustment member or the annular case relative to theother of the adjustment member or the annular case in response tothermal energy.

In a further non-limiting embodiment of the foregoing ring-shapedcompliant support, the gas turbine engine component is a blade outer airseal.

In a further non-limiting embodiment of any of the foregoing ring-shapedcompliant supports, the gas turbine engine component is secured to theadjustment member and is radially inside both the annular case and theadjustment member.

A method of moving a gas turbine engine component according to yetanother exemplary aspect of the present disclosure includes, among otherthings, exposing an annular case and an adjustment member to thermalenergy to turn the adjustment member relative to the annual case about acentral axis of a gas turbine engine, and supporting a gas turbineengine component with the adjustment member.

In a further non-limiting embodiment of the foregoing method, the gasturbine engine component comprises a blade outer air seal.

In a further non-limiting embodiment of either of the foregoing methods,the method includes expanding portions of the adjustment member againstan annular case to initiate rotation.

Although the different examples have the specific components shown inthe illustrations, embodiments of this disclosure are not limited tothose particular combinations. It is possible to use some of thecomponents or features from one of the examples in combination withfeatures or components from another one of the examples.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic, cross-sectional view of a gas turbine engine.

FIG. 2 is a cross-section of an area the gas turbine engine having acompliant support.

FIG. 3 is a highly schematic section view of the area at line 2-2 inFIG. 1.

FIG. 4 shows an exploded schematic view of the compliant support of FIG.3.

FIG. 5 is a section view at line 5-5 in FIG. 4.

FIG. 6 shows a portion of the compliant support at an initial position.

FIG. 7 shows an area 7 of the compliant support of FIG. 6.

FIG. 8 shows the FIG. 6 portion of the compliant support at a positionsubsequent to the initial position.

FIG. 9 shows the area 7 at the position subsequent to the initialposition.

FIG. 10 shows the FIG. 6 portion of the compliant support at a finalposition.

FIG. 11 shows the area 7 at the final position.

FIG. 12 is a section view of a second embodiment similar to FIG. 5.

FIG. 13 is a cross-section view of another embodiment of the compliantsupport of FIG. 4.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates an example gas turbine engine 20 thatincludes a fan section 22, a compressor section 24, a combustor section26 and a turbine section 28. Alternative engines might include anaugmenter section (not shown) among other systems or features. The fansection 22 drives air along a bypass flow path B while the compressorsection 24 draws air in along a core flow path C where air is compressedand communicated to a combustor section 26. In the combustor section 26,air is mixed with fuel and ignited to generate a high temperatureexhaust gas stream that expands through the turbine section 28 whereenergy is extracted and utilized to drive the fan section 22 and thecompressor section 24.

Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including direct drive and three-spool architectures.

The example engine 20 generally includes a low speed spool 30 and a highspeed spool 32 mounted for rotation about an engine central longitudinalaxis A relative to an engine static structure 36 via several bearingsystems 38. It should be understood that various bearing systems 38 atvarious locations may alternatively or additionally be provided.

The low speed spool 30 generally includes an inner shaft 40 thatconnects a fan 42 and a low pressure (or first) compressor section 44 toa low pressure (or first) turbine section 46. The inner shaft 40 drivesthe fan 42 through a speed change device, such as a geared architecture48, to drive the fan 42 at a lower speed than the low speed spool 30.The high-speed spool 32 includes an outer shaft 50 that interconnects ahigh pressure (or second) compressor section 52 and a high pressure (orsecond) turbine section 54. The inner shaft 40 and the outer shaft 50are concentric to one another and rotate via the bearing systems 38about the engine central longitudinal axis A.

A combustor 56 is arranged between the high pressure compressor 52 andthe high pressure turbine 54. In one example, the high pressure turbine54 includes at least two stages to provide a double stage high pressureturbine 54. In another example, the high pressure turbine 54 includesonly a single stage. As used herein, a “high pressure” compressor orturbine experiences a higher pressure than a corresponding “lowpressure” compressor or turbine.

The example low pressure turbine 46 has a pressure ratio that is greaterthan about five (5). The pressure ratio of the example low pressureturbine 46 is measured prior to an inlet of the low pressure turbine 46as related to the pressure measured at the outlet of the low pressureturbine 46 prior to an exhaust nozzle.

A mid-turbine frame 58 of the engine static structure 36 is arrangedgenerally between the high pressure turbine 54 and the low pressureturbine 46. The mid-turbine frame 58 further supports bearing systems 38in the turbine section 28 as well as setting airflow entering the lowpressure turbine 46.

The core airflow C is compressed by the low pressure compressor 44, thenby the high pressure compressor 52, then mixed with fuel and ignited inthe combustor 56 to produce high temperature exhaust gases that are thenexpanded through the high pressure turbine 54 and low pressure turbine46. The mid-turbine frame 58 includes vanes 60 that are in the coreairflow path and function as an inlet guide vane for the low pressureturbine 46. Utilizing the vane 60 of the mid-turbine frame 58 as theinlet guide vane for low pressure turbine 46 decreases the length of thelow pressure turbine 46 without increasing the axial length of themid-turbine frame 58. Reducing or eliminating the number of vanes in thelow pressure turbine 46 favorably shortens the axial length of theturbine section 28. Thus, the compactness of the gas turbine engine 20is increased and a higher power density may be achieved.

The disclosed gas turbine engine 20, in one example, is a high-bypassgeared aircraft engine. In a further example, the gas turbine engine 20includes a bypass ratio greater than about six (6:1), with an exampleembodiment being greater than about ten (10:1). The example gearedarchitecture 48 is an epicyclical gear train, such as a planetary gearsystem, star gear system or other known gear system, with a gearreduction ratio of greater than about 2.3.

In one disclosed embodiment, the gas turbine engine 20 includes a bypassratio greater than about ten (10:1) and the fan diameter issignificantly larger than an outer diameter of the low pressurecompressor 44. It should be understood, however, that the aboveparameters are only exemplary of one embodiment of a gas turbine engineincluding a geared architecture and that the present disclosure isapplicable to other gas turbine engines.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft., withthe engine at its best fuel consumption—also known as bucket cruiseThrust Specific Fuel Consumption (TSFC)—is the industry standardparameter of pound-mass (lbm) of fuel per hour being burned divided bypound-force (lbf) of thrust the engine produces at that minimum point.

“Low fan pressure ratio” is the pressure ratio across the fan bladealone, without a Fan Exit Guide Vane (“FEGV”) system. The low fanpressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.50. In another non-limiting embodimentthe low fan pressure ratio is less than about 1.45.

“Low corrected fan tip speed” is the actual fan tip speed in ft/secdivided by an industry standard temperature correction of [(Tram °R)/(518.7° R)]̂0.5. The “Low corrected fan tip speed”, as disclosedherein according to one non-limiting embodiment, is less than about 1150ft/second.

The example gas turbine engine includes the fan 42 that comprises in onenon-limiting embodiment less than about 26 fan blades. In anothernon-limiting embodiment, the fan section 22 includes less than about 20fan blades. Moreover, in one disclosed embodiment the low pressureturbine 46 includes no more than about 6 turbine rotors schematicallyindicated at 34. In another non-limiting example embodiment, the lowpressure turbine 46 includes about 3 turbine rotors. A ratio between thenumber of fan blades and the number of low pressure turbine rotors isbetween about 3.3 and about 8.6. The example low pressure turbine 46provides the driving power to rotate the fan section 22 and thereforethe relationship between the number of turbine rotors 34 in the lowpressure turbine 46 and the number of blades in the fan section 22disclose an example gas turbine engine 20 with increased power transferefficiency.

Referring now to FIGS. 2 and 3 with continuing reference to FIG. 1,components of the gas turbine engine 20 include a rotor disk 66, a blade68, vanes 70A, 70B, and a Blade Outer Air Seal (BOAS) 72. In thisexemplary embodiment, the components are part of the high pressureturbine 54. However, it should be understood that other portions of thegas turbine engine 20 could benefit from the teachings of thisdisclosure, including but not limited to, the compressor section 24 andthe low pressure turbine 46.

The high pressure turbine section 54 has a cross section 200. Duringoperation, area 210 of the cross-section may be hotter than area 220.Components that rotate about the axis, such as the rotor disk 66 and theblade 68, rotate through both the areas during operation. The rotatingcomponents experience relatively even wear.

If components do not rotate, the components may wear differentlydepending on their proximity to the relatively hot 220 or the less hotarea 210. The examples of this disclosure cause typically non-rotatingcomponents to rotate. This lessens variations in wear due to the area220 being hotter than the area 210.

In this exemplary embodiment, the rotor disk 66 (only one shown,although multiple disks could be used) is mounted to the outer shaft 50and rotates as a unit with respect to the engine static structure 36.Alternating rows of the rotating blades 68 are mounted to the rotor disk66. The vanes 70A and 70B of vane assemblies 70 are also supportedwithin an outer casing 74 of the engine static structure 36.

Each blade 68 includes a blade tip 68T that is positioned at a radiallyoutermost portion of the blades 68. The blade tip 68T extends toward theBOAS 72.

The BOAS 72 is disposed in an annulus radially between the outer casing74 and the blade tip 68T. The BOAS 72 generally includes a multitude ofBOAS segments 76 (only one shown in FIG. 2). The BOAS segments 76 mayform a full ring hoop assembly that encircles associated blades 68 of astage.

The BOAS 72, a type of gas turbine engine component, is held by aring-shaped compliant support assembly 78 that is mounted radiallyinward from the outer casing 74. The support 78 can include forward andaft flanges 80A, 80B that mountably receive the BOAS segments 76. Theforward flange 80A and the aft flange 80B may be manufactured of ametallic alloy material and may be circumferentially segmented for thereceipt of the BOAS segments 76.

The support 78 defines a cavity 82 that spans axially between theforward flange 80A and the aft flange 80B and radially between thesupport 78 and the BOAS segment 76. A secondary cooling airflow S may becommunicated into the cavity 82 to provide a dedicated source of coolingairflow for cooling the BOAS segments 76. The secondary cooling airflowS can be sourced from the high pressure compressor 52 or any otherupstream portion of the gas turbine engine 20.

Referring now to FIGS. 4-11 with continuing reference to FIGS. 1 and 2,the example support 78 includes an annular or ring-shaped case 84 and anadjustment member 88. The annular case 84 may form a portion of theouter casing 74. The annular case 84 defines a bore 90 that receives theadjustment member 88. The diameter of the bore 90 is closely matched tothe outer diameter of the adjustment member 88. In some examples, theadjustment member 88 is oversized relative to the bore 90 such that theadjustment member 88 is secured within the annular case 84 via aninterference fit. The tight fit provides a large frictional forcebetween the adjustment member 88 and the annular case 84. The adjustmentmember 88 may include the flanges 80A and 80B.

The example adjustment member 88 includes wings 94 extending radiallyand circumferentially from a ring 96. Slots 98 are defined within theadjustment member 88 to define the wings 94. The example slots 98 havean angle α from 30 to 75 degrees from a radial direction R when theengine 20 is at rest.

The adjustment member 88 includes a first array 100 of the wings 94 anda second array 104 of the wings 94. The first array 100 is axiallyupstream from the second array 104 relative to a direction of flowthrough the engine. The first array 100 and the second array 104 arearranged in an axially sequential manner The first array 100 is movablerelative to the second array 104.

In some examples, the first array 100 and the second array 104 aremachined from separate pieces of material. The rings 96 of the separatepieces are then secured directly to each other to form the adjustmentmember 88.

The adjustment member 88 is made of a first material. The annular case84 is made of a second material. The first material has a highcoefficient of thermal expansion relative to the second material.Example first materials include nickel alloy and steel.

During an operational cycle of the engine 20, the support 78 is exposedto elevated temperatures. The adjustment member 88 expands in responseto thermal energy levels more quickly than the ring 96. In this example,expansion of the adjustment member 88 causes the wings 94 to push, inthe direction D, against the annular case 84.

Expansion in the direction D is limited by the annular case 84. Relativeto the adjustment member 88, the annular case 84 expands very little, ifat all. Thus, the forces exerted in the direction D cause the adjustmentmember 88, and particularly the ring 96, to rotate in a direction Rrelative to the annular case 84. The forces exerted in the direction Dovercome any circumferential loading exerted on the adjustment member 88by the attached gas turbine engine component.

The adjustment member 88 rotates from an initial position shown in FIG.6. The BOAS segments 76 are attached to the ring 96 of the adjustmentmember 88. Rotating the adjustment member 88 thus rotates the BOASsegments 76 in the direction R. The rotation of the BOAS segments 76from the start position may be very small, such as less than amillimeter. In some examples, the amount of rotation during anacceleration and deceleration of the engine 20 is about 0.0127millimeters.

During cool down of the engine 20, the wings 94 contract. The firstarray 100 and the second array 104 expand and contract at differentrates to help prevent the adjustment member 88 and the BOAS segments 76from rotating back to the inital position during a cool down.

In this example, the first array 100 of wings 94 is axially thinner thanthe second array 104 of wings 94. During a thermal cycle, the thinnerwings 94 in the first array 100 will expand faster than the thickerwings 94 in the second array 104. The wings 94 in the first array 100will also contract during cool down, faster than the wings 94 in thesecond array 104. The wings 94 attached to the same circumferentiallocation of the ring 96 thus become circumferentially misaligned duringsome portion of a thermal cycle, which facilitates movement in thedirection R.

FIGS. 8 and 9 show the wings 94 during a cool down of the engine 20. Asshown, the wings 94 in the first array 100 are retracted more than thewing in the second array 104. The wings 94 that have retracted slip inthe direction R relative to the annular case 84. The wings 94 that haveretracted slip along the annular case 84 and interface with a differentsurface of the annular case 84. Notably, the wings 94 in the secondarray 104, which are relatively expanded, prevent the ring 96 fromreturning to the start position when the wings 94 in the first array 100contract. The amount of slippage corresponds generally to the differencein expansion between the wings 94 of the first array 100 and the wings94 of the second array 104.

After more cooling, the wings 94 in the second array 104 retract. Thewings 94 in the first array 100 prevent the ring 96 from returning tothe start position when the wings 94 in the first array 100 contract.

During a warm up of the engine 20, the wings 94 in the first array 100expand relative to the wings 94 in the second array 104. This causes thewings 94 in the second array 104 to slip relative to the wings 94 in thefirst array 100.

The wings 94 effectively walk, creep, or crawl the adjustment member 88in the direction R around the axis A to cause the BOAS 72 to turnrelative to the annular case 84. Thermal energy cycles of the engine 20are responsible for the rotation. After the successive thermal energycycles, the BOAS 72 may rotate further and further about the axis A. Inso doing, the circumferential position of the BOAS segment 76 within theengine 20 changes over time. Damage due to hot spots at particularcircumferential positions is thus lessened or eliminated.

In this example, the adjustment member 88 is radially inside the annularcase 84. In other examples, the adjustment member 88 could be radiallyoutside the annular case 84.

In this example, the BOAS 72, a type of gas turbine engine component, isattached to the adjustment member 88. In another example, the gasturbine engine component may be instead attached to the annular case 84.In such an example, exposure to thermal energy may turn the annular case84 rather than the adjustment member 88.

Referring to FIG. 12, in another example support 78 a, an adjustmentmember 88 a includes a first array 100 a of wings 94 a and a secondarray 104 a of wings 94 a. The wings 94 a in the first array 100 a andsecond array 104 a have the same axial thickness. The wings 94 a in thesecond array 104 a are coated with a thermally protective barriercoating 112. The coating 112 causes the wings 94 a in the second array104 a to heat and cool at a different rate that the wings in the firstarray 100 a.

In other examples, the wings are approximately the same size, and havethe same coefficient of expansion, one side, however, is exposed to theair stream and another is insulated. One array of the wings is thus moreexposed to a hotter area of the engine 20, than the other array. Themore exposed array of wings heats faster than the other array of wings.

Referring to FIG. 13, another example support 78 b is used to supportanother type of turbine engine component, such as a vane assembly 116.The support 78 b includes an annular case 84 b and an adjustment member88 b. An attachment member 120 secures a front end of the vane assembly116 to the adjustment member 88 b. A back end of the vane assembly 116is not directly secured to any fixed structure and follows the rotationsof the front end by the support 78 b.

Features of the disclosed examples include rotating a component using acompliant assembly. The component ordinarily having a fixedcircumferential position to avoid thermal energy damage. The compliantassembly may find beneficial use in many industries including aerospace,industrial, electricity generation, naval propulsion, pumps for gas andoil transmission, aircraft propulsion, vehicle engines and stationerypower plants.

The foregoing description shall be interpreted as illustrative and notin any limiting sense. A worker of ordinary skill in the art wouldrecognize that various modifications could come within the scope of thisdisclosure. For these reasons, the following claims should be studied todetermine the true scope and content of this disclosure.

I claim:
 1. A method of moving a gas turbine engine component,comprising; exposing an annular case and an adjustment member to thermalenergy to turn the adjustment member relative to an annual case about acentral axis of a gas turbine engine; and supporting a gas turbineengine component with the adjustment member.
 2. The method of claim 1,wherein the gas turbine engine component comprises a blade outer airseal.
 3. The method of claim 1, further comprising expanding portions ofthe adjustment member against an annular case to initiate rotation. 4.The method of claim 1, wherein the annular case defines a bore and theadjustment member is received within the bore.
 5. The method of claim 1,wherein the adjustment member comprises a first array of wings extendingradially and circumferentially from the ring, and a second array ofwings extending radially and circumferentially from the ring, whereinthe wings in the first array are axially thinner than the wings in thesecond array, and the wings in the first array are upstream from thewings in the second array.
 6. The method of claim 5, wherein slots inthe adjustment member separate the plurality of wings from each other.7. The method of claim 5, wherein the first array of wings are axiallysequentially aligned with the second array of wings.
 8. The method ofclaim 5, wherein the wings in the first array are insulated relative tothe wings in the second array, and the wings in the second array areuninsulated relative to the wings in the first array.
 9. A method ofmoving a gas turbine engine component, comprising: in response tothermal variations, elongating wings in a first and a second array ofwings to turn an adjustment member relative to an annular case, theadjustment member including the first array of wings extending radiallyand circumferentially from a ring, and the second array of wingsextending radially and circumferentially from the ring, the wings in thefirst array axially spaced from the wings in the second array.
 10. Themethod of claim 9, wherein the adjustment member and the annular casehave different rates of thermal expansion such that the adjustmentmember expands more quickly than the annular case in response to thermalenergy.
 11. The method of claim 9, wherein the rings in the first arrayare axially thinner than the rings in the second array.
 12. The methodof claim 9, wherein the rings in the first array are axially upstreamfrom the rings in the second array.
 13. A method of moving a gas turbineengine component, comprising: turning a gas turbine engine componentwith an adjustment member or an annular case relative to the other ofthe adjustment member or the annular case in response to thermal energy,the adjustment member having a ring, a first array of wings, and asecond array of wings that is axially spaced from the first array, thefirst and second array of wings extending radially and circumferentiallyfrom the ring.
 14. The method of claim 11, wherein the gas turbineengine component is a blade outer air seal.
 15. The method of claim 11,wherein the gas turbine engine component is secured to the adjustmentmember and is radially inside both the annular case and the adjustmentmember.
 16. The method of claim 11, wherein the wings in the first andsecond arrays elongate in response to thermal variations, the elongatingof the wings in the first array, the wings in the second array, or bothturning the adjustment member or the annular case relative to the otherof the adjustment member or the annular case.
 17. The method of claim11, wherein the wings in the first array are axially thinner than thewings in the second array.
 18. The method of claim 11, wherein the ringsin the first array are axially thinner than the rings in the secondarray.
 19. The method of claim 11, wherein the rings in the first arrayare axially upstream from the rings in the second array.